This invention relates generally to gas turbine engines, and in particular, to a cooled flow path surface region.
This application references co-pending applications assigned to the assignee of the present invention, which are identified as Attorney Docket No. 13DV-13513 and entitled xe2x80x9cDirectly Cooled Thermal Barrier Coating System,xe2x80x9d 13DV-13527 entitled xe2x80x9cMulti-layer Thermal Barrier Coating with Integrated Cooling System,xe2x80x9d the contents of which are incorporated herein by reference.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine, which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling air to the outer surfaces of the metal parts through various methods. The metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are the metal parts forming combustors and located aft of the combustors, including the turbine blades and vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate components downstream of the combustors, such as the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at metal surface temperatures of up to 2100xc2x0-2200xc2x0 F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved cooling designs and insulating thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air from the compressor directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a TBC is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts within engines to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component. TBC""s have also been used in combination with film cooling techniques wherein an array of fine holes extends from the hollow core through the TBC to bleed cooling air onto the outer surface of the TBC.
U.S. Pat. No. 6,039,537 to Scheurlen is directed to a turbine blade at least partly covered by a heat insulating layer system, wherein, the turbine blade has at least one interior space and a plurality of bores leading from the interior space out of the substrate. Some of the bores are covered by a heat-insulating layer, while others are not. The uncovered bores provide for film cooling. The heat insulating layer system is constructed such that flow of coolant through the covered bores is not necessary when the heat insulating layer system is intact. When a failure of the heat insulating layer system occurs, additional, previously covered bores open to aid in film cooling. Because there is no flow through the covered bores, there is no transpiration cooling.
U.S. Pat. No. 5,967,755 to Czech et al. also discloses bores passing through a covering layer having free outlet orifices. The bores remain open, not allowing for transpiration cooling.
There have also been attempts to allow cooling fluid to pass through a covering layer. For example, U. S. Pat. No. 3,240,468 to Watts et al. is directed to a hollow turbine blade jacketed with a permeable sheathing having a plurality of discrete recesses disposed behind the sheathing, to which recesses cooling fluid may be distributed through the hollow blade. The permeable sheathing is formed of powdered metal or ceramic pressed to the desired shape and dimensions and sintered. The cooling air passes through the skin of the blade forming a boundary layer around the blade. The sintering process will reduce the flow of cooling fluid available to form the boundary layer.
U.S. Pat. No. 4,067,662 to Rossmann is directed to a turbine engine blade whose outer shroud is constituted of a through-porous material. A cooling medium is able to flow through a passageway from the central core of the blade to permeate through the porous outer shroud up to the surface of the airfoil, where it forms a cooling boundary layer. The outer shroud of the core consists of a foamed, through-porous ceramic material, essentially constituted of aluminum oxide.
TBCs include well-known ceramic coatings, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBC""s usually do not adhere well directly to the superalloys used in the substrates. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the TBC. The bond coat may be made of a nickel-containing overlay alloy, such as a NiCrAIY or a NiCoCrAIY, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide whose surface oxidizes to a protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and overlying TBC are frequently referred to as a thermal barrier system.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. In some instances, the alloying elements of the bond coat can interdiffuse with the substrate alloy. Over time, as the airfoils are refurbished, walls of the airfoils are consumed, which reduces load carrying capability and limits blade life. This interdiffusion can also reduce the environmental resistance of the coating. Even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling, because reducing the demand for cooling contributes to improved overall engine operating efficiency.
While superalloys coated with thermal barrier coating systems do provide substantially improved performance over uncoated materials, there remains room for improvement. Film cooling is achieved by passing cooling air through discrete film cooling holes, typically having hole diameters ranging in size from about 0.015xe2x80x3 to about 0.030xe2x80x3. The film cooling holes are typically drilled with laser, EDM or ES machining. Due to mechanical limitations, each film hole has an angle ranging from 20xc2x0 to 90xc2x0 relative to the external surface. Therefore, each film jet exits from the hole with a velocity component perpendicular to the surface. Because of this vertical velocity component and a flow circulation around each jet due to the gas mixing, each film jet will have a tendency to lift or blow off from the external surface and mix with the hot exhaust gases, resulting in poor film cooling effectiveness.
Thus, there is an ongoing need for an improved thermal barrier coating system, wherein the environmental resistance and long-term stability of the thermal barrier coating system is improved so that higher engine efficiencies can be achieved. The bond coat temperature limit is critical to the TBC""s life and is kept lower than the temperature at the outer surface of the TBC, having an upper limit of about 2100xc2x0 F. Once the bond coat exceeds this temperature, the coating system can quickly deteriorate, due to high temperature mechanical deformation and oxidation, as well as increased interdiffusion of elements with the substrate alloy. The coating system can separate from the substrate exposing the underlying superalloy component to environmental damage from the hot gasses of combustion.
What is needed are improved designs that will allow turbine engine components to run at higher operating temperatures, thus improving engine performance without additional cooling air. It is also desirable to have a system that can take advantage of the thermal insulation provided by TBC. The present invention fulfills this need, and further provides related advantages.
The present invention provides for a method for transpiration cooling of the flow path surface region of an engine component used in a gas turbine engine comprising the steps of channeling a substrate to provide cooling channels or apertures through the substrate having a diameter to a diameter of about 0.0005xe2x80x3 to about 0.02xe2x80x3 to allow passage of cooling fluid from a cooling fluid source; applying a bond coat of about 0.0005xe2x80x3 to about 0.005xe2x80x3 in thickness to the substrate such that the bond coat partially fills the channels; applying a porous TBC of at least about 0.003xe2x80x3 to about 0.01xe2x80x3 thick to the bond coat, such that the TBC completely fills the channels; and, passing cooling fluid from a cooling fluid source through the channel and porous TBC. The cooling channels have a first end terminating as an exit orifice located on the surface of the substrate. The channel extends through the substrate and has a second end that is in fluid communication with a cooling circuit manufactured into the turbine engine component. The cooling circuit is connected to a cooling fluid supply that supplies air to the cooling fluid supply. Because the channel exit is filled with TBC, the cooling fluid is transmitted through the porous passageways of the TBC. The porous passageways are interconnected and provide a plurality of tortuous routes to the TBC surface.
The present invention further comprises both the cooled flow path surface region formed by the foregoing methods and the turbine component with the porous TBC for cooling the component.
In a different embodiment, the present invention comprises a cooling channel having a first and second end, the first end terminating in an exit orifice located on or in proximity to the surface of a substrate, the second end connecting to a cooling circuit contained within a turbine engine component. The cooling channel preferably has a diameter of about 0.002xe2x80x3 to about 0.008xe2x80x3. Applied to the substrate is a bond coat of about 0.0005xe2x80x3 to about 0.005xe2x80x3 in thickness, which partially fills the exit orifice, and first channel end. Applied to the bond coat is a porous TBC of at least about 0.003xe2x80x3 to about 0.01xe2x80x3 thick such that the porous TBC fills the remainder of the exit orifice and first channel end not filled by the bond coat.
An advantage of the present invention is the flow path surface region of the coated gas turbine component is actively cooled by transpiration cooling through the TBC. Transpiration cooling through the TBC lowers the TBC temperature, and allows a greater thermal gradient between the hot exhaust gas stream and the bond coat. By removing heat from this region, the integrity of the bond coat can be maintained at higher engine operating temperatures, resulting in a more efficient usage of cooling air to achieve a higher turbine engine efficiency and performance.
Another advantage of the present invention is that because the TBC is processed to be porous, cooling air is able to flow through TBC passageways and spread inside the TBC layer before exiting the TBC at its surface at a plurality of exit points at low volumes through transpiration cooling instead of exiting as discrete film jets at higher pressures and volumes, thereby preventing cooling film blow off from the TBC surface.
Still another advantage of the present invention is that the TBC filled holes have more flow resistance than open holes and, therefore, provide a more effective cross-sectional hole area compared to unfilled larger holes for flowing the same amount of air, resulting in more efficient heat transfer.
Yet another advantage is that the characteristics of many smaller passageways acting as holes with smaller flow cross-section areas will remove heat over a much larger heat transfer area more efficiently and will provide better cooling of the bond coat and the substrate than that provided by fewer larger holes expelling the same volume of cooling fluid.
Still another advantage of the present invention is that more effective cooling results in the reduction or elimination of sintering in the ceramic top coat at elevated engine operating temperatures, so that low conductivity is maintained in the TBC throughout its life.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying figures which illustrate, by way of example, the principles of the invention.